California State UniversityMFDC Lab.

Combustion- Propulsion TeamStudents: Amir Massoudi – Justin

Rencher

Andrew Clark – Uche Ofoma

Professor: Darrell Guillaume

Feb. 10, 2004

OBJECTIVESOBJECTIVES

Improve combustor performance in Ramjet and Scramjet engines by

optimizing air-fuel mixing to reduce pollutant formation and to increase

engine efficiency Validate the CFD Software called “Fluent”

• Verify that it can accurately predict the products of combustion

• Verify that is can accurately predict energy output

• Verify that it produces CFD result that are consistent with STARS

Model both Ramjet and Scramjet engines

• Modify fuel injection locations

• Alter the fuel-air ratios

• Modify the combustor geometry

Develop a Scramjet engine model that accurately predicts engine thrust

given parameters such as angle of attack, speed, and altitude Seek out all published data on Scramjet engines

Develop a Fluent model of a Scramjet

• Compare model performance to published data

• Run model under a variety of conditions to develop a look-up table to be used with the testbed.

AMIR MASSOUDI AMIR MASSOUDI (Graduate Student California (Graduate Student California State Univ. L.A)State Univ. L.A)

OBJECTIVE

• Build 2D combustion chamber model with numerical software

• Make a physical combustor based on results produced by computer models

• Compare results from two models

Equivalence Ratio & Equation of Burning Equivalence Ratio & Equation of Burning Hydrocarbon FuelHydrocarbon Fuel

AirRatioFuelmmFA af /

FuelRatioAirmmAF fa / Airofmassm

Fuelofmassm

a

f

ACTSTOICHSTOICHACT AFAFFAFA )/()()/()(

11

1

Running lean: Oxygen in exhaust

Running rich: CO and fuel in exhaust

Stoichiometric

General equation of combusting hydrocarbon fuel, excess air remaining after CO2 and H2O are formed

222222 21

79)1()

4(

221

79

4NO

yxOH

yCOxNO

yxHC yx

Geometry and Boundary Geometry and Boundary ConditionsConditions

Diameter 70 mm

Wall

Chamber Wall

300 mm

Interior

InteriorPressure Outlet

COMBUSTION CHAMBER DATA

• Fuel: n-Heptane (Gas and Liquid)

• Oxidizer: Air (%79 N2 - %21 O2)

• Vertical Chamber

• Parallel Injections for Fuel and Air (Study Velocity Inlet)

• Range between 0.55-0.95

CONTOURS OF STATIC TEMPERATURE (K) & MASS CONTOURS OF STATIC TEMPERATURE (K) & MASS FRACTION OF CO2 FOR GAS HEPTANEFRACTION OF CO2 FOR GAS HEPTANE

Static Temperature Mass Fraction CO2

CONTOURS OF STATIC TEMPRATURE (K) & MASS CONTOURS OF STATIC TEMPRATURE (K) & MASS FRACTION OF CO2 FOR LIQUID HEPTANEFRACTION OF CO2 FOR LIQUID HEPTANE

Static Temperature Mass Fraction CO2

ANDREW CLARKANDREW CLARK (Intern from Univ. of (Intern from Univ. of Manchester England)Manchester England)

Objectives• Find a non-technical method of creating a thermodynamic database for Fluent. This

would allow the usage of liquid aviation fuels which are not currently contained in

Fluent’s original thermodynamic database.

• Validate Fluent as a CFD code by comparing lift and drag coefficients obtained in

Fluent with coefficients obtained experimentally and coefficients obtained with

STARS.

Thermodynamic DatabaseThermodynamic Database

Summary of Fluent’s Thermodynamic Database:

• Contains NASA thermodynamic polynomials

• Thermodynamic polynomials are used to find thermodynamic and thermochemical properties

of species within a temperature range

• Thermodynamic database used primarily for combustion and propulsion.

• Fluent uses a modified CHEMKIN II Format database

• Database was created in MS Access and mail-merged to MS Word

Reasons to Construct Database:

• Update the current fuel types found in the Fluent database. More up-to-date polynomials can

be used, most of Fluent’s data is from the 1980’s where the source database is updated monthly

• Be able to utilize new fuel types.

Thermodynamic DatabaseThermodynamic Database

• NASA Thermodynamic polynomials have the form

45

34

2321 TaTaTaTaaR

CP

TaTaTaTaTaaRTH

64

53

42

321 5432

64

53

42

321 432ln aTaTaTaTaTaRS

• Completed thermodynamic tables for three fuels

(n-Heptane gas, n-Heptane liquid, Jet A liquid)

• Used data from Caltech

• Fluent has different format for Polynomial Coefficients

• Converted polynomial coefficients from source format to Fluent format

Validating FluentValidating Fluent

• 2D Subsonic validation using JavaFoil (panel method) to produce theoretical data for a NACA 4415 airfoil

• 2D Supersonic validation using linearised supersonic airfoil theory for a diamond airfoil

• 3D Subsonic validation using STARS data supplied by CFD team for Titan (a NASA award winning student design)

• 3D Supersonic and Hypersonic validation using NASA’s report for a Winged-Cone GHV and the CFD team’s results from a CFD research code called STARS

Fluent Results were Compared to:Fluent Results were Compared to:

ResultsResults

2D Subsonic

Validation

Successful – Spalart-Allmaras Turbulence Model was found to

produce the most accurate results.

2D Supersonic

Validation

Successful – Inviscid Solver was found to produce the most

accurate results.

3D Supersonic

Validation

Successful - Inviscid Solver was found to produce the most

accurate results.

3D Hypersonic

Validation

Successful - Inviscid Solver was found to produce the most

accurate results.

A Lift Coefficient Comparison between NASA's, CFD Team's and Fluent's Computational Results for the Winged Cone GHV at Mach 4

-0.05

0

0.05

0.1

0.15

0.2

0 2 4 6 8 10 12

Angle of Attack (Degrees)

Lif

t C

oeff

icie

nt

NASA Report S-A Solver CFD Team Inviscid Solver

3D Dimensional Supersonic Validation of Winged Cone 3D Dimensional Supersonic Validation of Winged Cone GHV At Mach 4GHV At Mach 4

A Drag Coefficient Comparison between NASA's, CFD Team's and Fluent's Computational Results for the Wing Cone GHV at Mach 4

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0 2 4 6 8 10 12

Angle of Attack (Degrees)

Dra

g C

oef

fici

ent

S-A Solver CFD Team Inviscid Solver NASA Report

3D Dimensional Supersonic Validation of Winged Cone 3D Dimensional Supersonic Validation of Winged Cone GHV At Mach 4GHV At Mach 4

Uche OfomaUche Ofoma (Graduate Student California State (Graduate Student California State

Univ. L.A)Univ. L.A) Objective

• Seek out all published results on Scramjet engines

Results Many tests have been performed at NASA Langley Results are classified so we cannot get them

Other Engine Performance Data (Tunnel)Other Engine Performance Data (Tunnel)

Results from 2001 CIAM tunnel tests

Gaseous hydrogen used as fuel

Mach 6 flow velocity

Approx. 75 kg thrust measured

Other Engine Performance Data (Tunnel)Other Engine Performance Data (Tunnel)

Published data from NASA/CIAM Hypersonic Flying

Laboratory (Feb. 1998)

Other Engine Performance Data (Tunnel)Other Engine Performance Data (Tunnel)

Japan’s NAL Kaduka Space

Propulsion Laboratory scramjet

engine tests at Mach 4, 6 and 8

Tests similar to NASA Langley’s

Net thrust of 500 N produced

Engine ModelEngine Model

Analyze NASA Langley, CIAM,

NAL, etc. scramjet test data for

performance curves, altitude, fuel

consumption, speed, flight angle

of attack, emissions, etc.

Compare test data to Fluent

Model

Create engine analysis

methodology for use as a design

tool (spreadsheet or program

code)

Output engine data will provide

results for CFD team

AltitudeSpeed

Pitch Change

Input engine parameters

Performance look-up tables

Output engine parameters

Angle of attack AltitudeSpeed

Input engine parameters

Performance look-up tables

Output engine parameters

Angle of attack AltitudeSpeed

Justin RencherJustin Rencher (Undergraduate Student California (Undergraduate Student California

State Univ. L.A)State Univ. L.A) Objectives

• Accurately simulate supersonic combustion of an appropriate fuel in a two dimensional scramjet using the CFD software, Fluent.

Approach

• Build geometry and cases based upon existing research and results, applying known methods and accepted approaches to the Fluent CFD environment.

Results

• Building supersonic combusting ramjet simulations within Fluent that actually converge has proven to be quite difficult.

• Information on how to create such simulations is scarce and sometimes classified. • Observations made at the recent AIAA conference in Reno have shown us that we are on the

right track.

Description of Current TaskDescription of Current Task

• The geometry for this particular model is based on published data from the

NASA Langley Scramjet Test Complex

• The focus of these CFD cases is primarily on the behavior of fluid flow and combustion

characteristics as they are affected by what are known as ramp injectors.

• These ramp injectors are utilized to enhance fuel/air mixing so that the combustor length can

be reduced.

• A ramp angle of 10.3 deg was used in published data. The following slides show results of 10.3,

12.3, and 8.3 deg angles as determined by Fluent CFD Software.

• Results of airflow and combustion for each model are pictured. A Mach 2 airflow is used and

combustion is carried out with gaseous n-heptane.

X/G=16G (gap length) = 3 in

Shock Wave Diagram with Shock Wave Diagram with Ramp InjectorRamp Injector

Combustor Duct

X/G = 16G (gap length) = 3 in

8.3 deg Ramp Angle: Mach 2 Airflow and Combustion8.3 deg Ramp Angle: Mach 2 Airflow and Combustion

• The two top slides are air flow only, displaying contours of mach number for the 8.3 degree ramp injectors

• The slide to the left displays contours of static temperature for air flow with combustion

10.3 deg Ramp Angle: Mach 2 Airflow With No Combustion

10.3 deg Ramp Angle: Mach 2 Airflow and Combustion10.3 deg Ramp Angle: Mach 2 Airflow and Combustion

• The two top slides are air flow only, displaying contours of mach number for the 10.3 degree ramp injectors

• The slide to the left displays contours of static temperature for air flow with combustion

12.3 deg Ramp Angle: Mach 2 Airflow 12.3 deg Ramp Angle: Mach 2 Airflow and Combustionand Combustion

• The two top slides are air flow only, displaying contours of mach number for the 12.3 degree ramp injectors

• The slide to the left displays contours of static temperature for air flow with combustion